1. Field of the Invention
The present invention is in the field of helicopters, and more specifically relates to a structure for the tail rotor of a helicopter.
2. The Prior Art
The blades of the helicopter tail rotor are normally connected to the tail rotor shaft in such a manner as to accommodate several degrees of freedom of motion of the rotor blade. These types of motion include: pitch, by which the angle of attack of the blade is changed; flap, the component of the motion of the blade tips parallel to the tail rotor shaft axis; and lead-lag, the motion of the blades in the plane of the rotor with respect to the rotor shaft.
These motions were accommodated in earlier helicopters through the provision of hinges. The use of hinges increased the mechanical complexity of the rotor structure, imposed lubrication requirements, and was not as efficient on a strength-to-weight basis as subsequently-developed techniques. Typically, four individual blades were separately attached to the tail rotor through the hinges, and thus, the centrifugal loads were borne by the rotor hub and the hinge mechanisms. Alternatively, for small helicopters, blade-pair assemblies were attached to the hub by pivot bearings, thus containing the centrifugal loads within the blade-pair assembly, but still requiring bearings at the hinge points.
The development of anisotropic composite structural elements including those in which high-strength fibers are embedded in an epoxy matrix, has enabled development of hingeless rotors. Typically in such a rotor, the diametrically opposed blades are connected by an uninterrupted beam of composite material having its fibers extending radially to withstand the centrifugal forces and to provide low feathering stiffness. Typically, the beam is relatively thin in the direction of the rotor shaft and relatively broad in the chordwise direction. This cross-sectional shape facilitates flapping of the blades while permitting the driving torque to be transmitted to the blades, and blade pitch control is effected by elastically twisting the beam. Because the motions of the blades are accommodated by elastic flexing of the beam, the beam is usually referred to as a flexbeam. In some rotors, the flexbeam extends radially only a relatively short distance, and the aerodynamic portion of the blade is attached to the end of the flexbeam.
Because the flexbeam is rigidly affixed to the rotor hub, twisting of the flexbeam for pitch control requires the application of the twist-producing force at a location spaced radially outward from the hub, and spaced forward or aft of the blade's twist axis in the chordwise direction. Typically, the force is transmitted through a control push rod, the end of which is pivotally connected to a pitch horn affixed to a torsionally-stiff pitch case (torque tube) which surrounds the flexbeam but is spaced from it and which extends radially outwardly to the aerodynamic portion of the blade to which is it attached and to which it can transmit torque. With regard to flapping motion, the flexbeam may be though of as possessing an effective flapping hinge about which the flapping motion of the blade is centered. Normally, the pitch horn is located radially outwardly of the effective flapping hinge on the leading edge side of the blade, so that when the blade flaps upwardly, the blade pitch angle decreases (see FIG. 11). Alternatively, the same effect may be had by locating the pitch horn radially inboard of the effective flapping hinge on the trailing edge side of the blade. Either of these locations of the pitch horn results in reduced blade thrust and strap stresses with controlled flapping motion of the rotor blade. The magnitude of the angle .delta..sub.3 is indicative of the amount of pitch-flap coupling, where .delta..sub.3 is the complement of the angle between the radial direction and the line through the effective flapping hinge and the point on the pitch horn at which the pitch control forces are applied.
The advantages of hingeless rotors are not obtained without introducing other complications, including the potential for aeroelastic instabilities peculiar to the hingeless rotors. These instabilities were identified in the early 1970's, and various solutions were tried, as will be described below.
For example, in U.S. Pat. No. 3,999,886 issued Dec. 28, 1976 to Ormiston et al., it was disclosed that the stability can be improved by inclining the principal elastic flexural axes and by including an arrangement for varying the pitch of the blade in relation to the degree of bending of the blade in a plane parallel to the plane of rotation of the blade, i.e., for providing pitch-lag coupling. In one embodiment, the pitch-lag coupling is provided by arranging the parts in such a manner that the torque tube forces the blade shank to twist in response to bending of the blade shank in a direction parallel to the plane of rotation. In another embodiment, pitch-lag coupling is produced by proper positioning of the pitch link to force the blade shank to twist when lead-lag bending of the blade shank occurs.
In the helicopter rotor disclosed by Ormiston et al., the centrifugal forces of the opposed blade halves are applied to the hub, which requires that the hub be sufficiently strong to withstand these stresses. It is not possible to join the opposed blades by a single flexbeam extending through the hub because of the anisoelastic structure of the flexbeam disclosed. In that structure, strips of high stiffness and of low stiffness are combined so that the principal elastic axes are inclined to the geometric axes of the cross section of the flexbeam. If a beam of this structure extended on both sides of the rotor, the principal elastic axes would be correctly aligned on one side of the hub, but would be incorrectly aligned on the opposite side of the hub. This prevents the flexbeam from extending through the hub and accordingly requires that the centrifugal stresses be borne by the hub.
The advantages of extending the flexbeam through the hub were recognized by workers at United Technology Corporation in Hartford, Conn., and their discoveries are disclosed in the following publications: U.S. Pat. No. 4,047,839 issued Sept. 13, 1977 to Ferris et al.; the technical paper "Composite Bearingless Tail Rotor for UTTAS" by Fenaughty and Noehren, presented by the 32nd Annual National Forum of the American Helicopter Society, May, 1976; the technical paper "Aeroelastic Characteristics of Composite Bearingless Rotor Blades" by Richard L. Bielawa, presented at the 32nd Annual National V/SPOL Forum of the American Helicopter Society, Washington, D.C., May, 1976; and U.S. Pat. No. 4,087,203, issued May 2, 1978 to Ferris.
Broadly, these references disclose a helicopter rotor having opposed blades interconnected by a common flexible spar which passes across the rotor axis and is connected to the drive shaft by clamped hub plates, as shown in FIG. 1 herein. A spanwise-extending torque tube having a pitch horn at its radially inner end forms a rigid connection with the spar and blade at the radially outwardly end of the torque tube.
In the United Technology Corporation's approach, the flexbeam was tapered but was relatively thick at the hub, and was bolted rigidly to the hub as best seen in FIG. 3 of U.S. Pat. No. 4,047,839. This construction caused loads to be transmitted through the hub which had a tendency to crack as a result. As described in U.S. Pat. No. 4,087,203, an elastomeric snubber was used to center the pitch case with respect to the flexbeam because of control considerations. The snubber induced a large flap bending moment, and this had a tendency to crack the relatively thick rib that was provided at the radially inward end of the blade where the pitch case and flexbeam joined the blade. As will be seen below, the structure of the present invention permits these disadvantages to be avoided.
Still another approach to the design of a bearingless tail rotor is disclosed in the technical paper "The YUH-61A Tail Rotor: Development of a Stiff Inplane Bearingless Flexstrap Design", by John Shaw, Jr. and W. Thomas Edwards of the Boeing Vertol Company, presented at the 33rd Annual National Forum of the Helicopter Society, Washington, D.C., May, 1977. As illustrated in FIG. 2 herein, a relatively thin, untapered flexbeam is used. The flexbeam is attached to the hub by torsionally-soft "cross-flexures", the purpose of which is to eliminate clamping of the flexstrap by the hub. The blade is bolted to the flexbeam, and this requires reinforcing the ends of the flexbeam. To avoid large flapping angles and corresponding flexbeam stresses, a large .delta..sub.3 angle of -65.degree. is used. Consistent with standard nomenclature, a negative value of .delta..sub.3 indicates "down" pitch for "up" flap. This effectively induces a positive pitch-lag kinematic coupling (with the pitch links inclined from the vertical between the swash plate and pitch horn) which has an adverse flap-lag instability, as described in the above paper. To cure this instability, the blades had to be swept back. Also, this design requires a relatively large motion of the pitch link.
Thus, it is seen that although the technology of flexbeam rotor design was relatively advanced at the time of the present invention, nevertheless, a number of problems were inherent in the approaches then used, and it is an object of the present invention to overcome these problems.